Optimized Cross-Ply Orientation in Composite Laminates
A composite laminate has a primary axis of loading and comprises a plurality resin plies each reinforced with unidirectional fibers. The laminate includes cross-plies with fiber orientations optimized to resist bending and torsional loads along the primary axis of loading.
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1. Field:
The present disclosure generally relates to composite laminates, such as those used in aircraft, and deals more particularly with a fiber reinforced resin laminate having cross-plies with optimized fiber orientations.
2. Background:
Fiber reinforced resin laminates, such as carbon fiber reinforced plastics (CFRP), are widely used in aerospace and other applications because of their favorable strength-to-weight ratio. These composite laminates may be fabricated by laying up multiple plies of unidirectional reinforcing fibers held in a resin matrix, commonly known as prepreg. The plies in the layup may have differing fiber orientations arranged in an order that results in the required laminate strength and stiffness for a particular application. For example, in aircraft skins, the composite laminate may comprise groups of plies respectively having reinforcing fibers oriented at 0, +45, −45 and 90 degrees relative to a reference axis, with the majority of the plies being +/− degree plies. Although the number of plies in the laminate may vary at different locations along the wing, the angular orientation of the plies may be substantially constant over the length of the wing, and therefore not fully optimized to match performance requirements at individual locations on the wing. The use of these differing ply orientations allow the composite laminate structure to better resist bending, shear, torsional and bearing loads for a given application, but the use of constant orientations of the plies may result in a skin that is heavier than desired for a particular application.
Reducing the weight of composite laminate structures used in aircraft applications may improve the operating efficiency of an aircraft. The weight of such structures may depend, at least in part, on the number of plies in the laminate, which in turn may be determined by the strength and stiffness requirements for the particular application.
Accordingly, there is a need for a composite laminate that employs a reduced number of total plies while retaining the required laminate strength, rigidity and resistance to splitting and crack propagation. There is also a need for a composite laminate aircraft skin that optimizes the balance between the skin's bending and torsional strength and stiffness by using cross-plies that have optimized fiber orientations. Further, there is a need for a composite laminate skin exhibiting reduced weight through variation of the ply orientation over the wing length.
SUMMARYThe disclosed embodiments provide a composite laminate structure, such as an aircraft skin, that has reduced weight and may improve the laminate's structural strength, stiffness and resistance to splitting and crack propagation. The weight of the structure is reduced by using fewer 0 degree plies in the composite laminate. Reduction in the number of 0 degree plies in the composite laminate may also reduce labor and material costs. The reduction in the number of 0 degree plies is achieved without substantially reducing the composite laminate's shear properties, including torsional strength and stiffness, by optimizing the orientation of the cross-plies. Cross-ply orientation is optimized by more closely lining up the fiber orientations of the cross-plies with the primary longitudinal loads on the composite laminate structure. The benefits from the disclosed cross-ply optimization may be implemented in aircraft skins designs without substantial architectural changes in the wing design. A 3 to 5 percent reduction in the weight of a wingbox may be achieved using the disclosed ply optimization technique.
In some applications, the cross-plies may have fiber orientations within a range of approximately 33 and degrees, while in other applications, a cross-ply orientation within the range approximately 23 and 45 degrees may provide optimum composite laminate performance while achieving weight reduction. The disclosed composite laminates may be employed as skins on wings, stabilizers, control services or other structures used in aircraft. In a wing skin application of the disclosed composite laminate, for example and without limitation, it may be possible to achieve a three to five reduction in the weight of an aircraft wing box while increasing the bending performance of the wing box. The optimized orientation angle of the cross-plies may constant or variable over an area of a composite laminate. For example, and without limitation, the optimized orientation angle of the cross-plies may be variable in a span-wise direction along the length of a wing skin.
According to one disclosed embodiment, a composite laminate comprises at least one resin ply reinforced with unidirectional fibers having a substantially 0 degree fiber orientation relative to a reference axis, and at least one resin ply reinforced with unidirectional fibers having a substantially 90 degree fiber orientation relative to the reference axis. The composite laminate further comprises cross-plies of resin reinforced with unidirectional fibers each having ±θ degree fiber orientations relative to the reference axis, where θ is within the range of approximately 25 degrees and 43 degrees. In some applications, θ may be within the range of approximately 35 degrees and 40 degrees, and may vary in angular orientation over a length or within an area of the laminate.
According to another disclosed embodiment, a composite laminate aircraft skin is provided having a primary axis of loading. The aircraft skin comprises a first group of fiber reinforced resin plies having a fiber orientation substantially parallel to the axis of loading, a second group of fiber reinforced resin plies having a fiber orientation substantially perpendicular to the axis of loading, and a third group of fiber reinforced resin plies having a ±θ degree fiber orientation relative to the axis of loading, where θ is within the range of approximately 33 and 43 degrees. In some applications, θ may be within the range of approximately 35 and 40 degrees.
According to still another embodiment, a composite laminate aircraft skin having a primary axis of loading comprises at least one group of fiber reinforced resin cross-plies having fiber orientations that vary along the primary axis of loading. The fiber orientations of the cross-plies may vary within a range of approximately 25 and 45 degrees, and in some applications, the skin may further include cross-plies having fiber orientations of approximately 45 degrees relative to the axis of primary loading.
According to a further embodiment, a method is provided of manufacturing a composite laminate having a primary axis of loading. The method comprises assembling a multi-ply layup, including laying up a first set of resin plies each reinforced with unidirectional fibers having ±θ degree fiber orientations relative to the axis of loading, where θ is within the range of approximately 25 degrees and 43 degrees, laying up a second set of resin plies each reinforced with unidirectional fibers having a substantially 0 degree fiber orientation relative the primary axis of loading, and laying up a third set of resin plies each reinforced with unidirectional fibers having a substantially 90 degree fiber orientation relative the primary axis of loading. The method further includes laminating the first, second, and third sets of plies together. θ is within the range of approximately 35 degrees and 45 degrees. Laying up the first set of resin plies includes varying the fiber orientation θ along the primary axis of loading. Laying up the first set of resin plies and varying the fiber orientation may be performed using a computer controlled automatic fiber placement machine. The method may further comprise drilling at least one hole through the laminated plies, and/or cutting through at least an edge of the laminated plies.
According to still another embodiment, a method is provided of fabricating a composite aircraft wing skin having a primary axis of loading. The method comprises laying a plurality of resin plies reinforced with unidirectional fibers, including orienting a first set of the plies generally parallel to the primary axis of loading, and orienting a second set of the plies at ±θ degree fiber orientation relative to the primary axis of loading, where θ is within the range of approximately 25 degrees and 43 degrees. In some applications, θ may be is within the range of approximately 35 degrees and 45 degrees. Orienting the second set of plies includes varying the angular orientation of the second of plies along the length of the primary axis of loading. Varying the angular orientation of the second set of plies includes placing plies in the second set thereof in a first angular orientation along a first stretch, and placing plies in the second set in second angular orientations along a second stretch that are different from the first angular orientation. The first angular orientation may be substantially constant along the first stretch, and the second angular orientations may vary along the second stretch.
The features, functions, and advantages can be achieved independently in various embodiments of the present disclosure or may be combined in yet other embodiments in which further details can be seen with reference to the following description and drawings.
The novel features believed characteristic of the advantageous embodiments are set forth in the appended claims. The advantageous embodiments, however, as well as a preferred mode of use, further objectives and advantages thereof, will best be understood by reference to the following detailed description of an advantageous embodiment of the present disclosure when read in conjunction with the accompanying drawings, wherein:
The disclosed embodiments relate to a composite laminate and related fabrication method that may be employed to fabricate any of a variety of composite laminate structures. The embodiments may find use in numerous fields, particularly in the transportation industry, including for example, aerospace, marine, automotive applications and other applications where light weight composite laminates are employed. Thus, referring now to
Each of the processes of method 20 may be performed or carried out by a system integrator, a third party, and/or an operator (e.g., a customer). For the purposes of this description, a system integrator may include without limitation any number of aircraft manufacturers and major-system subcontractors; a third party may include without limitation any number of vendors, subcontractors, and suppliers; and an operator may be an airline, leasing company, military entity, service organization, and so on.
As shown in
Systems and methods embodied herein may be employed during any one or more of the stages of the production and service method 20. For example, components or subassemblies corresponding to production process 28 may be fabricated or manufactured in a manner similar to components or subassemblies produced while the aircraft 22 is in service. Also, one or more apparatus embodiments, method embodiments, or a combination thereof may be utilized during the production stages 28 and 30, for example, by substantially expediting assembly of or reducing the cost of an aircraft 22. Composite laminate structures manufactured according to the disclosed embodiments may increase the strength and stiffness of components of the aircraft 22 while reducing the aircraft weight. Similarly, one or more of apparatus embodiments, method embodiments, or a combination thereof may be utilized while the aircraft 22 is in service, for example and without limitation, to maintenance and service 36.
The disclosed laminate 60 may be combined with other structures to form a part 62 such as the composite sandwich panel 76 shown in
Attention is now directed to
Ply 64b includes unidirectional reinforcing fibers 66b having angular fiber orientations +θ relative to the X axis, while ply 64c includes unidirectional reinforcing fibers 66c having angular fiber orientations −θ relative to the X axis. The plies 64b and 64c having fiber orientations of +θ and −θ respectively may sometimes also be referred to herein as “cross-plies” having angular orientations of ±θ, and ±θ may sometimes be referred to as the “cross-ply angle”. As will be discussed below, the particular cross-ply angle ±θ is optimized to maintain or improve the performance of the composite laminate 60 while reducing its weight. In some applications, the cross-ply angle ±θ may be within a range of approximately 10 and 43 degrees, while in other applications, desired results may be obtained where the cross-ply angle ±θ is within a range of approximately 33 and 43 degrees. In still other applications, use of a cross-ply angle ±θ within a range of approximately 35 and 40 degrees may provide beneficial or useful results.
The cross-ply angle ±θ may vary in magnitude over one or more areas of the part 62 (
For simplicity of illustration, only four plies 64a-64d are shown in the example of
By employing cross-plies 64b, 64c having optimized fiber orientations of +θ, −θ respectively, fewer zero degree plies 64d may be required for a particular application, such as a wing skin. Fewer 0 degree plies may be needed because the fiber orientations of the cross-plies 64b, 64c more closely line up with the primary axis of loading, i.e. the X axis, compared to conventionally used ±45 degree plies, thereby contributing to the bending strength and stiffness of the laminate 60 while maintaining the required level of torsional strength and stiffness. A small loss of torsional strength and stiffness resulting from the disclosed cross-ply optimization technique may not be particularly detrimental in most wing skin applications because the skins are designed with relatively large margins for torsional strength and stiffness.
A typical wing skin laminate may comprise 30/60/10 percent of 0, 45 and 90 degree plies, respectively. Since the majority of the plies may be 45 degree plies, it may be appreciated that optimizing the angle of the cross-plies may result in the need for fewer 0 degree plies 64d. As a result of the use of fewer 0 degree plies 64d, the weight of the composite laminate 60 may be reduced in those applications where most of the composite load resistance is in the 90 degree direction, which in the illustrated example is substantially parallel to the Y axis. Additionally, the use of cross-plies having angular fiber orientations ±θ may boost the bearing strength of the 0 degree plies 64d while assisting in the suppression or delay of potential splitting and/or crack propagation in the 0 degree plies 64d and the 90 degree plies 64a since the fibers 66b, 66c of the cross-plies 64b, 64c cross over and tie together the fibers 66a, 66d of the 0 degree plies 64d and the 90 degree plies 64a, respectively. The ability of the cross-plies 64b, 64c to suppress or delay ply splitting and crack propagation may be particularly important where holes 70 (
Attention is now directed to
At 100, a second set of fiber reinforced resin plies 64d is laid up, wherein each of the plies 64d has a generally 0 degree fiber orientation relative to the primary axis of loading. At 102, a third set of fiber reinforced resin plies 64a is laid up, wherein each of the plies 64a has a generally 90 degree fiber orientation relative to the primary axis of loading. At 104, the plies of the layup are laminated together by consolidating and curing the layup. At 106, optionally, the cured part may be trimmed to final size by cutting one or more edges 74 (
By optimizing the orientation angle ±θ of the cross-plies in the laminate 60 forming the skin 110, the bending strength and stiffness of the wing 38 and the skin 110 may be increased. Due to this increase in bending strength and stiffness, fewer 0 degree plies may be used in the laminate 60, resulting in a corresponding decrease in the weight of the skin 110 and thus of the wing 38. In other words, because some of the plies are better oriented to resist the main load paths, fewer plies are required that are oriented in a cross direction to the main load paths. This optimization of the cross-ply orientation angle ±θ results in a weight reduction of skin 110. Also, optimization of the cross ply orientation angle ±θ allows the wing skin 38 to be tailored at different points or stretches to better match local requirements to resist bending forces 118 and torsional forces 120. While a wing 38 is shown in
As previously mentioned, in some applications, it may be possible to vary the cross-ply angle ±θ over one or more local areas of the laminate 60 in order to optimize the local or overall performance of the laminate and/or reduce its weight. For example,
The description of the various embodiments has been presented for purposes of illustration and description, and is not intended to be exhaustive or limited to the embodiments in the form disclosed. Many modifications and variations will be apparent to those of ordinary skill in the art. Further, different embodiments may provide different advantages as compared to other embodiments. The embodiment or embodiments selected are chosen and described in order to best explain the principles of the embodiments, the practical application, and to enable others of ordinary skill in the art to understand the disclosure for various embodiments with various modifications as are suited to the particular use contemplated.
Claims
1. A composite laminate having a primary axis of loading, comprising:
- a plurality of resin plies each reinforced with unidirectional fibers, and including cross-plies have fiber orientations optimized to resist bending and torsional loads along the primary axis of loading.
2. The composite laminate of claim 1, wherein the plurality of resin plies includes:
- at least one resin ply reinforced with unidirectional fibers having a substantially 0 degree fiber orientation relative to the primary axis of loading,
- at least one resin ply reinforced with unidirectional fibers having a substantially 90 degree fiber orientation relative to the primary axis of loading, and
- the cross-plies each having ±θ degree fiber orientations relative to the primary axis of loading, where θ is within a range of approximately 10 degrees and 43 degrees.
3. The composite laminate of claim 2, wherein θ is within a range of approximately 33 degrees and 43 degrees.
4. The composite laminate of claim 2, wherein θ is within a range of approximately 35 degrees and 40 degrees.
5. The composite laminate of claim 2, wherein the θ degree fiber orientations of the cross-plies vary along the primary axis of loading.
6. The composite laminate of claim 2, wherein the θ degree fiber orientations of the cross-plies vary within an area of the composite laminate.
7. The composite laminate of claim 2, wherein the θ degree fiber orientations of the cross-plies are selected to match loading of composite laminate along the primary axis.
8. A composite laminate aircraft skin having a primary axis of loading, comprising:
- a first group of fiber reinforced resin plies having a fiber orientation substantially parallel to the primary axis of loading;
- a second group of fiber reinforced resin plies having a fiber orientation substantially perpendicular to the primary axis of loading; and
- a third group of fiber reinforced resin cross-plies extending traverse to the plies in the first and second groups thereof and having a ±θ degree fiber orientation relative to the primary axis of loading, where θ is optimized to resist bending loads and torsional loads applied to the skin.
9. The composite laminate aircraft skin of claim 8, wherein θ is within a range of approximately 33 and 43 degrees.
10. The composite laminate aircraft skin of claim 8, wherein θ is within a range of approximately 35 and 40 degrees.
11. The composite laminate aircraft skin of claim 8, wherein θ varies across the skin along the primary axis of loading.
12. The composite laminate aircraft skin of claim 8 wherein:
- the skin is a wing skin having a wing root and a wing tip,
- θ is approximately 43 degrees in the area of the wing root, and
- θ is approximately 10 degrees at the wing tip.
13. A composite laminate aircraft skin having a primary axis of loading, comprising:
- at least one group of fiber reinforced resin cross-plies having fiber orientations that vary along the primary axis of loading.
14. The composite laminate aircraft skin of claim 13, wherein the fiber orientations of the cross-plies vary within a range of approximately 10 and 43 degrees.
15. The composite laminate aircraft skin of claim 13, further comprising:
- plies having fiber orientations of approximately 90 degrees relative to the primary axis of loading, and
- plies having fiber orientations of approximately 0 degrees relative to the primary axis of loading.
16. A method of laying up a composite aircraft skin having a primary axis of loading, comprising:
- laying up cross-plies of resin reinforced with unidirectional fibers, including orienting the cross-plies at angles that vary along the primary axis of loading.
17. The method of claim 16, wherein the angles are within the range of approximately +10 and +43 degrees, and −10 and −43 degrees relative to the primary axis of loading.
18. The method of claim 16, wherein orienting the cross-plies includes selecting orientation angles for the cross plies that are substantially matched to loads imposed on the wing skin at multiple locations along the primary axis of loading.
19. A method of fabricating a composite aircraft wing skin having a primary axis of loading, comprising:
- laying up a plurality of resin plies each reinforced with unidirectional fibers having a fiber orientation substantially parallel to the primary axis of loading;
- laying up a plurality of resin plies each reinforced with unidirectional fibers having a fiber orientation substantially orthogonal to the primary axis of loading; and
- laying up a plurality of resin cross-plies each reinforced with unidirectional fibers and having an angular fiber orientation ±θ, including optimizing the angular fiber orientation ±θ to substantially match the loading on the wing skin.
20. The method of claim 19, wherein optimizing the angular orientation ±θ is performed at each of a plurality of locations along the primary axis of loading.
21. The method of claim 19, wherein optimizing the angular orientation ±θ includes selecting a fiber orientation angle within the range of approximately 10 and 43 degrees.
22. A method of reducing the weight of a composite wing skin formed of fiber reinforced plies that include 0 degree and 90 degree plies traversed by cross-plies, comprising:
- reducing the number of 0 degree plies required to resist loads on the wing skin by optimizing the angular orientation of the cross plies.
23. The method of claim 22, wherein optimizing the angular orientation of the cross-plies includes selecting an angular orientation ±θ for the cross-plies, where ±θ is within the range of approximately 10 and 43 degrees.
24. The method of claim 22, wherein the 90 degree plies extend substantially parallel to a primary axis of loading of the wing skin, and optimizing the angular orientation of the cross plies is performed at more than one location in the span-wise direction of the wing skin along the primary axis of loading.
Type: Application
Filed: Jun 8, 2012
Publication Date: Dec 12, 2013
Patent Grant number: 9289949
Applicant: THE BOEING COMPANY (Chicago, IL)
Inventor: Max U. Kismarton (Renton, WA)
Application Number: 13/491,784
International Classification: B32B 5/12 (20060101); B32B 37/00 (20060101);